Method of making and repairing turbine blades

ABSTRACT

The turbine blade tip clearance control system presently used with existing turbine blade uses an abradable seal material with a conventional squealer tip. Others use a thin, coating added to the tip of the blade. The present turbine blade tip control system minimizes the problem of controlling the clearance between the blade and the shroud and overcomes the problem of wearing away of the thin coating added to the tip by using a rub tolerant, high-temperature seal material coating on a turbine shroud and a turbine blade made of at least two materials. The combination of the coating and an outer tip being of a different material than the metallic body also minimizes the problems associated with the burning of dirty fuels. The blade has a metallic body of high strength and the outer tip has a strength less than the strength of the body and made of a material resistant to oxidation, sulfidation and thermal fatigue at operating temperatures. The outer tip has a radial length &#34;L&#34; at least 1 mm in length.

This is a continuation of application No. 07/377,495, filed July 10,1989 now abandoned, which is a division of application No. 07/236,874filed Aug. 26, 1988 now abandoned.

DESCRIPTION

1. Technical Field

This invention relates generally to gas turbine engine blade tipclearance control and more particularly to improvements for increasingthe turbine blade and shroud assembly resistance to high temperature,oxidation, corrosion and sulfidation.

2. Background Art

Gas turbine engines and other turbomachines have rows of wheel mountedblades which rotate within a generally cylindrical case or shroud. Suchengines are generally driven by directing high temperature gastherethrough to cause the blades to rotate relative to the shroud. Thegas is generally corrosive in nature due to the chemical makeup thereof.The blades may be coated with a thin protective coating to protect themfrom the corrosive action of the gas. The purpose of the shroud is toprevent gas from bypassing the blades. Without the shroud, the gas couldflow outwardly of the radially outer end, or tip, of the blade. Tominimize the amount of gas escaping between the tip of a rotor blade andthe shroud, the operating clearance between the tip of the rotor bladeand the shroud should be as small as is practical. Generally the lengthof the blade is selected such that the radially outer end, or tip, ofthe blade is disposed close enough to the inner surface of the shroud soas to form a seal therebetween. One of the problems encountered withsuch engines is that rubbing contact inevitably occurs between the bladetip and the shroud. The main cause of such rubbing is difference inthermal expansion and/or contraction between the turbine and the shroud.The immediate problem caused by the rubbing is that the blade and theshroud wear away resulting in eventual loss of efficiency. The farreaching effect is that the protective coating at the tip of the bladeis worn away exposing more of the base metal to the corrosive gases andmore rapid deterioration of the blade tip will occur.

Some users may want to burn low grade ashforming fuels which containhigh amounts of corrosive elements. Price, availability, and flexibilityof fuel supply requires the consideration of crude and residual oil forfuture industrial gas turbine installations. Other fuels such as low Btugas from enhanced oil recovery, fire flood operations or otherindustrial processes are available. Marine operations, for example,occasionally inadvertently contaminate fuel with sea water. Thesecriteria have led turbine manufacturers to develop technologies forsuccessful gas turbine operation on such low grade or "dirty" fuels.

One of the major concerns in using lower grade fuels is the corrosioncaused by various mineral elements contained in these fuels. Corrosionrates rapidly increase at metal temperatures above 650° C. Thetemperature at the exposed turbine tip exceeds this temperature. Thus,actions must be taken to overcome the corrosion problems which erode theouter portion of turbine blades and the inwardly facing surface ofshrouds affixed about the turbine blades.

The use of a coating of material on the radially outer edges of a bladetip has been suggested as a way of protecting the tip from the gases.However, due to the physical characteristics of present day coatingssuch coatings normally have a thickness of between 5 and 30 mils andwill eventually be worn away. Since each such rub wears away some of thematerial, the radially thicker the coating, the more rubs it willwithstand before it is completely worn away. If the coating is not ofsufficient thickness to withstand these rubs the base blade material isexposed to the corrosive gases and rapid corrosion will occur. However,there is a maximum usable thickness limitation of the coating due to thelack of structural rigidity of the coating compared to the relativelyhigh structural rigidity of the remainder of the tip. That is, if thecoating was too thick radially, relative to the radial length of theblade, one rub could cause the entire coating of material to break off.Another problem resulting from the wearing away of the blade tip resultsin a greater gap of space between the blade and the shroud. The effectsof this problem will be explained later. When the coating is provided ona superalloy turbine blade tip, the method of application must bemetallurgically compatible with the superalloy substrate so that theproperties of the substrate are not degraded. Such considerations placerestraints on the kinds of coatings and processing techniques which areuseful in the fabrication of such coatings.

Furthermore, structural integrity of the blades are absolutely essentialdue to the high centrifugal stresses and elevated temperatures to whichthe blades are exposed during high-speed rotation under normal operatingconditions.

Examples of the above structures are described in the following patents:U.S. Pat. No. 4,390,320 issued to James E. Eiswerth on June 28, 1983,U.S. Pat. No. 4,689,242 issued to Roscoe A. Pike on Aug. 25, 1987, U.S.Pat. No. 4,589,823 issued to William K. Koffel on May 20, 1986, and U.S.Pat. No. 4,610,698 issued to Harry E. Eaton, et al on Sept. 9, 1986.

Other attempts to resist corrosion of the blade tip have resulted invarious combinations of blade tips. For example, U.S. Pat. No. 4,232,995issued to Kenneth W. Stalker, et al an Nov. 11, 1980, discloses an innerand outer tip portion bonded to a metallic projection body and the innertip portion respectively. The inner tip is diffusion bonded to the body.

An example of rebuilding a portion of a blade disclosed in U.S. Pat. No.3,574,924 issued to Gordon L. Dibble on Apr. 13, 1971. In thatrebuilding process the damaged area of the blade is trimmed off andreplaced with a precisely correspondingly sized replacement portion. Thereplacement portion is diffusion bonded in a mold to exactly duplicatethe contour of the original blade. Another method for repairing bladetips is disclosed in U.S. Pat. No. 4,214,355 issued to John W. Zelahy onJuly 29, 1980. A first member is bonded to the side walls of a hollowbody and a second member is bonded to the first member. Both members arediffusion bonded to the side walls and each other.

The coating patents listed above fail to provide a satisfactory tip onthe blade which will withstand a multitude of rubs. The physicalcharacteristics of todays coatings prevent the functional use of athickness which will withstand a multitude of rubs which occurs withtodays engines. The internal bond of todays coatings will allow thematerial to shear or break off from the blade body. Thus, heretofore acoating having a thickness greater than between 5 and 30 mils is notworkable.

In todays applications attempts to increase turbine blade and shroudassembly resistance to high temperature, oxidation, corrosion andsulfidation have failed to consider that rubs are inevitable and thatthey must compensate for the affect that they have on the blade tip andshroud assembly.

The continued requirement by industry for use of dirty fuels has causedattempts to minimize the affects of these fuels on the turbinecomponents. For example, the coating patents and the replacement portionor new portion patents listed above employ a diffusion bonding techniquewhich uses heat and pressure to achieve the atomic bond therebetween.Furthermore, the disclosures in such above-mentioned patents do notconsider the length required of a tip portion to insure a quantity ofmaterial resistant to oxidation and sulfidation that will remain after amultitude of rubs have occurred.

The greater the gap or spacing between the blades and the shroud thelower the efficiency of the engine. Thus, it is essential to provide assmall a gap as is practical to insure maximum efficiency of the engine.In the art as shown above, as the blades wear the gap becomes greaterand the efficiency is decreased. Therefor, it is desirable to preventwear of the blades.

The present invention is directed to overcome one or more of theproblems as set forth above.

DISCLOSURE OF THE INVENTION

In one aspect of the present invention an axial flow turbine for usewith a gas turbine engine is disclosed. The engine includes a rotatableturbine wheel having an annular array of turbine blades attachedthereto, a shroud assembly which has a radially inwardly facing surfacepositioned adjacent and radially outwardly of the turbine blades, and acoating affixed to the surface. The coating is resistant to hightemperature, oxidation, corrosion and sulfidation and is rub-tolerant.Each of the blades has a metallic body of a predetermined strengthsufficient to resist structural deformation at normal operatingparameters of the engine and an outer tip is attached to the body andpositioned adjacent the coating. The outer tip is comprised of amaterial which has a predetermined strength less than that of the bodyyet sufficient to prevent breakage thereof and separation from themetallic body during rubbing contact between the tip and the coating.The material composition of the outer tip is resistant to hightemperature, oxidation, corrosion, sulfidation and thermal fatigue atnormal operating parameters of the engine. The outer tip is formed onthe metallic body by a weld layered puddling build-up using a filler rodhaving the material as an ingredient thereof.

In another aspect of the invention a method of making a turbine bladehaving a metallic body of a predetermined strength to resist structuraldeformation and an outer tip formed on the metallic body. The methodcomprises the following steps. Forming the outer tip on an outerextremity of the metallic body by a weld layered puddling build-up usinga filler rod comprised of a material having a predetermined strengthless than that of the body and being resistant to high temperature,oxidation, corrosion, sulfidation and thermal fatigue at operationtemperatures. Machining the outer tip to a preestablished profile afterforming of the outer tip on the metallic body.

In another aspect of the present invention a method of repairing aturbine blade comprises the following steps. Removing material from ametallic body the material being removed forming an outer extremity ofthe metallic body. Forming the outer tip on an outer extremity of themetallic body by a weld layered puddling build-up using a filler rodcomprised of a material having a predetermined strength less than thatof the body and being resistant to high temperature, oxidation,corrosion, sulfidation and thermal fatigue at operating temperatures.Machining the outer tip to a preestablished profile after forming of theouter tip on the metallic body.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial sectional side view of a gas turbine enginedisclosing the turbine blade tip control system of this invention.

FIG. 2 is an enlarged sectional view of the turbine tip clearance areaencircled by line II of FIG. 1.

FIG. 3 is an enlarged sectional view taken along line III--III of FIG.2.

BEST MODE FOR CARRYING OUT THE INVENTION

Referring to FIG. 1, a gas turbine engine 10, not shown in its entirety,has been sectioned to show a turbine tip control system 12 forcontrolling internal leakage of a turbine section 14 of the engine. Theengine 10 includes an outer case 16, a combustor section 18, acompressor section 20 and a compressor discharge plenum 22 fluidlyconnected to the compressor section 20. The plenum 22 is partiallydefined by the outer case 16 and a multipiece inner wall 24 partiallysurrounding the turbine section 14 and the combustor section 18. Thecompressor section 20 includes a plurality of rotatable blades 26attached to a longitudinally extending center shaft 28 driven by agasifier turbine section 29. A plurality of compressor stator blades 30extend from the outer case 16 and are positioned between the rotatableblades. For illustration convenience, only a single stage of amultistage axial compressor 20 is shown. The combustor section 18includes a combustion chamber 32 positioned in the plenum 22 and aplurality of fuel nozzles 34 (one shown) positioned in the chamber 32 atthe end near the compressor section 20. The turbine section includes afirst stage turbine 36 disposed partially within an integral first stagenozzle and shroud assembly 38. The first stage turbine 36 includes arotatable turbine wheel 39 and an annular array of turbine blades 40(one shown) attached thereto. The assembly 38 is attached to the innerwall 24 and supported from the center shaft 28 by a bearing arrangement41 and a series of thermally varied masses 46 which are assembled toprevent rapid thermal variation during heating and cooling of suchmasses 46. A nozzle support case 48 is disposed within the outer case 16and is attached to the case 16 by a plurality of bolts and dowels notshown. An integral second stage nozzle and shroud assembly 50 isattached to the nozzle support case 48 by a plurality of nozzle hooks 52and a second stage turbine 54 is disposed partially within the assembly50. An integral third stage nozzle and shroud assembly 56 is attached tothe nozzle support case 48 by a plurality of nozzle hooks 58 and a thirdstage turbine 62 is disposed partially within the assembly 56. All theturbines 36, 54, 62 are connected to the longitudinally extending centershaft 28.

As more clearly shown in FIG. 2, the shroud assembly 38 has a radiallyinwardly facing surface 64 positioned adjacent and radially outwardly ofthe turbine blade 40. A coating 66 is affixed to the surface 64 and isresistant to high temperature, oxidation, corrosion and sulfidation andis rub-tolerant. Rub-tolerant means that the coating will separate fromitself in fine particle form or smear when rubbed by the blades ratherthan to break in larger pieces. The rub-tolerant coating is applied by aplasma spray process and has a thickness of approximately 0.7 mm. Thecoating is comprised of cobalt, nickel, chromium and aluminum. Thecoating has the physical characteristics and is applied in a process toresult in a uniform, continuous surface free from spalling, cracking,chipping or flaking.

Each of the turbine blades 40 has an attaching base 67 connected to theturbine wheel 39 and a blade portion 68 attached to the base 67. Theblade portion includes a metallic body 69 attached to the base 67 and anouter tip 70 formed outwardly on the body. The outer tip 70 has an outerface 71 thereon opposite the end formed on the body 69. The metallicbody 69 has a root area 72 near the base 67, an outer extremity 73outwardly of the root area 72 and a perimeter 74 around the body. Thebody 69 is of a material which has a predetermined strength sufficientto resist structural deformation at normal operating parameters of theengine 10. The outer extremity 73 is positioned adjacent the coating 66to provide a normal operating clearance therebetween. The outer tip 70can be formed to include a flat tip portion, not shown, or a squealertip portion 76. The squealer tip portion 76 provides a better sealbetween the blade 40 and the coating 66 than does a flat tip portion.The squealer tip portion 76 is formed to have a ridge 78 positioned atand around the perimeter 74 of the metallic body 69 and a concaveportion 80 is formed in the center of the ridge 78. The ridge 78 of thesquealer tip portion 76 has a tip extremity 82 and a bottom portion 84.As best shown in FIG. 3 the thickness of the ridge 78 is of a thinnercross-sectional area near the tip extremity 82 than is the thickness ofa cross-sectional area of the ridge 78 near the bottom portion 84resulting in the squealer tip portion 76 including a tapered build-upbeing thinner near the tip extremity 82 than near the bottom portion 84of the squealer tip portion 76. The above structure is a result of thewelding process since the layered build-up results in a narrower passper respective build-up pass or series of repetitive passes. Thus, thesquealer tip portion 76 is approximately 50% shorter than a conventionalsquealer tip configuration shown by phantom line 86. The outer tip 70 iscomprised of a material which has a predetermined strength less thanthat of the body 69 yet has sufficient strength to prevent breakingand/or separating from the metallic body 69 during hard rubbing contactsbetween the blade 40 and the coating 66 and/or the shroud 38. Thus, theouter tip 70 must have sufficient length to accept the wearing away of aportion of the material and yet not be totally removed from protectingthe metallic body 69 from oxidation, sulfidation and corrosion. Forexample, the metallic body 69 has a tensile strength of approximately1080 kPa at room temperature and approximately 750 kPa at 870° C. and ayield strength of approximately 930 kPa at room temperature andapproximately 650 kPa at 870° C. The outer tip 70 has a tensile strengthof approximately 870 kPa at room temperature and approximately 380 kPaat 870° C. and a yield strength of approximately 380 kPa at roomtemperature and approximately 230 kPa at 870° C. Furthermore, the outertip 70 is made of a nickel base material and is selected to includeapproximately 20%-24% chromium, 13%-15% tungsten, 1%-3% molybdenum and atrace of boron, lanthanum, silicon and manganese. A weld layeredpuddling build-up process is used to form the outer tip 70 on themetallic body 69. The process uses a filler rod having the nickel basematerial as described above as an ingredient thereof.

As best shown in FIG. 3, the outer tip 70 has a radial length "L"sufficient to insure protection of the metallic body 69 from thecorrosive gases at the highest operating temperature even after aplurality of rubs occur between the blade 40 and the coating 66 and/orthe shroud 38. Thus, the length of the outer tip 70 is at least 1 mm toinsure that if a portion of the tip 70 is worn away during the rubbingof the blade 40 and the coating 66 and/or shroud 38 a portion of thematerial remains on the metallic body 69.

INDUSTRIAL APPLICABILITY

The turbine tip control system 12 of the present invention is part of agas turbine engine 10. The compressor section 20 provides combustion airfor the engine 10. The air is mixed with fuel from the fuel nozzles 34in the combustion chamber 32. The mixture of fuel and air is combustedand the resulting gas expands and is used to drive the turbines 36, 54,62 The gas is directed into the first stage turbine 36 through the firststage nozzle-and shroud assembly 38. The nozzle guides the gas so thatit strikes the blades 40 at a preestablished angle to exert a maximumforce for rotating the turbine 36. Any leakage between the turbineblades 40 and the shroud assembly 38 results in reduced power and a lossof efficiency. Thus, the tightness or clearance between the outerextremity 73 of the turbine blade 40 and the shroud assembly 38 controlsefficiency. For example, the clearance between the turbine 36, 54, 62and the shroud assembly 38, 50, 56 in a typical application isapproximately 0.5 mm.

The coating 66 minimizes the affects of corrosion, oxidation andsulfidation on the turbine components caused by burning low grade fuels.The primary function of the coating is to prevent damage/wear of theblade tips by "sacrificing" itself locally wherever rubs occur. Thus,the tip of the blade is not worn away during normal rubs and theefficiency of the engine remains high since the clearance between theblade 40 and the coating 66 becomes greater at point of the rub ratherthan along the entire radius of the rotating blade 40. The coating hasthe ability to accept rubs without breaking The rubs will result in thecoating being displaced or ground away in a granular form, rather thanin pieces or chunks which could cause further damage to the engine. Itis desirable that the coating be displaced or ground away rather thanallowed to build-up on one surface or the other since the build-up wouldlead to even heavier rubs.

The turbine blade 40 has been made of at least two different materialsto overcome the high temperature, oxidation, corrosion and sulfidationproblems. The metallic body 69 has a predetermined strength to resistthe operating parameters of the engine. The engine of this embodimentoperates up to about 12,000 rpm which converts to a linear speed ofapproximately 475 m/s at the outer extremity 73 of the turbine blade 40and has a normal operating temperature range of between about 760° C. to930° C. Due to the mass of the blade, heat of the expanded gas and thecentrifugal force resulting from the high speed of rotation, thesestrengths as mentioned earlier are required to insure proper operationof the turbine blade 40. Stresses induced by temperature and speed aremost severe about the root area 72 of the turbine blade 40. For example,the greater mass near the root area 72 absorbs and stores more heat thanthe mass of the blade near the outer tip 70. Thus, the metallic body 69is made of a material having high strength but is more susceptible tocorrosion and oxidation. The outer tip 69 has a lower mass resulting inless heat absorption and lesser centrifugal force which provides anopportunity to use a material having lesser strength. Thus, the outertip 70 of the blade which is less susceptible to these extreme loads ismade of a material having lower strength but being more resistant tocorrosion, sulfidation and oxidation. For example, the material used inthis application has a predetermined strength less than that of themetallic body 69. The length of the outer tip 70 is of a sufficientlength to insure that a portion of the material remains on the metallicbody even after hard rubs occur. For example, the engine 10 is stopped,the shroud assembly 38, 50, 56 has cooled down, the turbine 36, 54, 62has remained hot and the engine 10 is restarted. Thus, the shroudassembly 38, 50, 56 is at a relative small diameter due to thermalcontraction and the turbine 36, 54, 62 is at a relative large diameterdue to the retention of heat and thermal expansion resulting in maximuminterference or a hard rub between the blade 40 and the coating 66and/or shroud 38. The outer tip 70 is formed on the metallic body 69 bya weld layered puddling build-up using a filler rod having the nickelbase material as the main ingredient thereof. After welding, the turbineblade 40 is finish machined by a grinding process to blend the weldedmaterial or outer tip 70 to match the airfoil contour of the metallicbody 69 so that the surfaces blend in a smooth and continuous manneraround the profile of the blade tip area. The concave portion 80 of thesquealer tip portion 76 remains as welded. The squealer tip portion 76although being shorter than that of a conventional squealer tip portionwill provide a better seal than that of the flat tip portion. Theshorter squealer tip portion 76 as compared to a conventional squealertip configuration 86 has little affect on the sealing between the blade40 and the coating 66 and/or shroud 38. Although the unmachined concaveportion 80 adds imbalance to the blade, the shorter height of thesquealer tip portion helps to reduce the amount of imbalance created bythe forming process. Turbine blades 40 which are not going to be coatedare heat treated in a conventional manner.

When repairing a turbine blade 40 which has a protective coatingresistant to oxidation and sulfidation applied to the entire blade, thecoating must be removed or stripped away before repair can begin.Material is removed at the outer end of the metallic body 69 bymachining, such as grinding. The quantity of material which was removedwill be replaced by a welding build-up process as discussed above. Theblending operation is preformed as discussed above. At this time, forblades which are to be coated by a conventional process do not require aseparate heat treating process since the coating process requires thatthe coating be cured in an inert atmosphere at approximately 760 ° for 4hours and the separate heat treating process is not required. The entireblade is normally coated due to economics but only the metallic body 69needs to be coated to resist oxidation, corrosion and sulfidation.

Other aspects, objects and advantages will become apparent from a studyof the specification, drawings and appended claims.

We claims:
 1. A method of making a turbine blade having a metallic bodybeing of a predetermined strength sufficient to resist structuraldeformation and an outer tip formed on the metallic body comprising thesteps of:(a) forming the outer tip on an outer extremity of the metallicbody by a weld layered puddling build-up of a repetitive series ofpasses using a filler rod comprised of a material having a predeterminedstrength less than that of the metallic body and being resistant to hightemperature, oxidation, corrosion, sulfidation and thermal fatigue atoperating temperatures; and (b) machining the outer tip to apreestablished profile after forming of the outer tip on the metallicbody.
 2. The method of making a turbine blade of claim 1 wherein thestep of forming the outer tip to the outer extremity of the metallicbody by the weld layered puddling build-up includes forming a squealertip portion.
 3. The method of making a turbine blade of claim 2 whereinthe step of forming the outer tip to the outer extremity of the metallicbody by the weld layered puddling build-up includes forming a squealertip portion approximately 50 percent of the length of a conventionalsquealer tip configuration.
 4. The method of making a turbine blade ofclaim 1 wherein the step of machining the outer tip to a preestablishedcontour includes a grinding process.
 5. The method of making a turbineblade of claim 1 wherein the method further includes a heat treatingprocess of the blade.
 6. The method of making a turbine blade of claim 1the method further includes a coating process on at least the metallicbody of the blade.
 7. The method of making a turbine blade of claim 6wherein the coating process further includes a heat treating process. 8.The method of making a turbine blade of claim 7 wherein the step of heattreating includes heat treating in an inert atmosphere at approximately760° C. for 4 hours.
 9. A method of repairing a turbine blade comprisingthe steps of:(a) removing material from a metallic body said removing ofmaterial resulting in an outer extremity being formed on the metallicbody; (b) forming an outer tip on the outer extremity of the metallicbody by a weld layered puddling build-up of a repetitive series ofpasses using a filler rod comprised of a material having a predeterminedstrength less than that of the body and being resistant to hightemperature, oxidation, corrosion, sulfidation and thermal fatigue atoperating temperatures; and (c) machining the outer tip to apreestablished profile after forming or the outer tip on the metallicbody.
 10. The method of repairing a turbine blade of claim 9 wherein thestep of removing material includes stripping a coating from the blade.11. The method of repairing a turbine blade of claim 9 wherein the stepof forming the outer tip to the outer extremity of the metallic body bythe weld layered puddling build-up includes forming a squealer tipportion.
 12. The method of repairing a turbine blade of claim 9 whereinthe step of forming the outer tip to the outer extremity of the metallicbody by the weld layered puddling build-up includes forming a squealertip portion having a height being approximately 50 percent that of aconventional squealer tip configuration.
 13. The method of repairing aturbine blade of claim 9 wherein the step of machining the outer tip toa preestablished contour includes a grinding process.
 14. The method ofrepairing a turbine blade of claim 9 the method further includes a heattreating process of the blade.
 15. The method of repairing a turbineblade of claim 9 wherein the method further includes a coating processon at least the metallic body of the blade.
 16. The method of repairinga turbine blade of claim 15 wherein the coating process further includesa heat treating process.
 17. The method of repairing a turbine blade ofclaim 16 wherein the coating process includes a process in which theblade is cured in an inert atmosphere at approximately 760° C. for 4hours.